A gas turbine engine typically includes a fan section, a compressor section, a combustor section, and a turbine section. Air entering the compressor section is compressed and delivered into the combustion section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section.
The turbine section includes at least one array of vanes axially offset from at least one array of blades. In order to increase operating efficiency of the gas turbine engine, operating temperatures in the turbine section have typically been increased. The elevated temperature of the turbine section requires cooling of several of the components in the turbine section with air from the compressor section. However, to meet the increased temperature demands in a gas turbine engine, the additional cooling that is required will increase the demand on the compressor section, which supplies the cooling fluid, reducing the overall efficiency of the gas turbine engine.
In one exemplary embodiment, a rotor blade assembly includes at least one rotor blade including an airfoil that has a leading edge internal cooling passage that extends through the rotor blade and is in communication with cooling holes along a leading edge of the airfoil. A compression portion includes a compression passage that is in communication with the leading edge internal cooling passage.
In a further embodiment of the above, at least one rotor blade includes a trialing edge internal cooling passage.
In a further embodiment of any of the above, the compression passage includes an inlet that has a first cross-sectional area and an outlet that has a second cross-sectional area smaller than the first cross sectional area.
In a further embodiment of any of the above, the leading edge internal cooling passage includes an aerodynamic shape.
In a further embodiment of any of the above, the aerodynamic shape tapers in a radially outward direction.
In a further embodiment of any of the above, the compression portion includes a compressor wheel forming a ring.
In a further embodiment of any of the above, the compression portion is integral with at least one rotor blade and includes a scoop.
In a further embodiment of any of the above, the leading edge internal cooling passage is configured to receive cooling air from an intermediate stage of a compressor section to provide bleed air to the leading edge internal cooling passage.
In another exemplary embodiment, a gas turbine engine assembly includes a compressor section for providing a bleed air source. A rotor blade assembly includes at least one rotor blade that includes an airfoil that has a leading edge internal cooling passage that extends through the rotor blade and is in communication with cooling holes along a leading edge of the airfoil. A compression portion includes a compression passage that is in communication with the leading edge internal cooling passage and the bleed air source.
In a further embodiment of any of the above, the bleed air source is from an intermediate stage of the compressor section.
In a further embodiment of any of the above, the compressor section includes a high pressure compressor and the intermediate stage is located in the high pressure compressor.
In a further embodiment of any of the above, the rotor blade includes a trialing edge internal cooling passage.
In a further embodiment of any of the above, the compression passage includes an inlet having a first cross-sectional area and an outlet having a second cross-sectional area smaller than the first cross sectional area.
In a further embodiment of any of the above, the leading edge internal cooling passage includes an aerodynamic shape.
In a further embodiment of any of the above, the aerodynamic shape tapers in a radially outward direction.
In a further embodiment of any of the above, the compression portion includes a compressor wheel forming a ring.
In a further embodiment of any of the above, the compression portion is integral with the at least one rotor blade and includes a scoop.
In another exemplary embodiment, a method of cooling a rotor blade includes directing bleed air from a compressor section into a rotor cavity. A first portion of the bleed air at a first pressure is directed into a compression portion for increasing a pressure of the bleed air entering a leading edge internal cooling passage of at least one rotor blade to a second pressure. A second portion of the bleed air at the first pressure is directed into a trailing edge internal cooling passage.
In a further embodiment of any of the above, the method includes rotating the compression portion with the at least one rotor blade.
In a further embodiment of any of the above, the method includes extracting the bleed air from an intermediate stage of a high pressure compressor in the compressor section.